Note: Descriptions are shown in the official language in which they were submitted.
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AIRCRAFT GA8 TU~IN~ ~NGINE BLE~D AIR
~:N~5RGY ~ECOV~ Y APPA~ATU8
BACRG}tOUND OF ~ CNVENq!ION
Field o~ the InventiQn
The invention relates to aircraft environmental control
systems powered by gas turbine engines for supplying
aircraft system bleed air. In particular, the invention
relates to an aircraft gas turbine engine powered bleed air
supply system for an aircraft environmental control system
wherein the ~xcess pressure energy is returned to the engine
to improve fuel efficiency and the system is also used to
reduce drag on engine narelles or other sur~aces by pumping
boundary layer air.
Description_of Related Art
Environmental control systems, commonly referred to as
ECS systems, incorporate various piaces o~ eguipment such as
turbocompressors, regulatinq valves, heat exchangers, and
other apparatus including what is referred to as an ECS pack
to condition engine bleed air. Modern day jet aircraft use
turbocomprQssors in the ECS packs of their environmental
controls systems to condition bleed air for use in the cabin
wherein the turboco~pressors ~re powered by the same bleed
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that is conditioned for cabin refreshing air and which is
usually supplied by the gas turbine engines which provide
aircraft propulsion. Other EcS systems employ auxiliary
power units (APU's) that are separate gas turbine engines,
not use for propulsion, to power ECS packs as well as other
aircraft equipment.
Bleed air is conventionally taken from the engine
compressor at a stage downstream of the variable vane
compressor stages so as not to interf~re with the operation
of the variable vane stages which greatly enhance the
efficiency of the gas turbine engine and ~reatly reduces the
specific fuel consumption (SFC) of the engine. The
compressor bleed air is cooled ~y fan air in a heat
exchanger conventionally referred to as a precooler and is
then delivered to the environmental con~rol system for
controlling cabin air freshness, pressure, and temperature.
The ECS conventionally includes two or more ECs packs
mounted in ECS bays on different sides of the aircraft which
receive compressor bleed air from the engines. The bleed
after being used to power the ECS pack and refresh the cabin
is then dumped overboard. All the energy remaining in the
bleed air dumped overboard cost fuel and therefore
represents a significant loss in specific fuel consumption.
Extraction of aircraft bleed air ~rom the engine
compressor has adverse affects on the propulsion cycle and
engine life. En~ine turbine power is needed to compress air
and account for compressor inefficiency. There~ore~ extra
fuel consumption is always associated with gas turbine
engine compressor bleed air (air which does not produce
thrust). This extra fuel burned in the engine combustor
results in higher gas temperature delivered to the engine
turbine and reduction o~ turbine blade life. Such penalties
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must be incurred in order that the engine turbine provide
extra power ascociated with bleed air.
It is no~ possible, without undue complexity, to always
bleed the engine compressor stage which provides exactly the
correct pressure needed for the aircraft anti-ice and ECS
systems. Typically, only two bleed ports are provided.
Therefore, the result is to bleed air which exceeds minimum
pressure requirements resulting in even higher pPnalty to
the engine cycle than would be required by the aircraft
systems.
Most often the bleed air is not only at a higher than
required pressure, it is also too hot. For reasons of fire
safety, maximum bleed air temperature is usually ~imited to
350 to 500F. Temperature control requires cooling thP
bleed air with a precooler. Most modern engines use fan air
to cool compressor bleed air. Use of fan air imposes an
additional penalty on fuel consumption. Further, the
precooler is usually large and requires a fan air scoop
which produces drag. A typical larse turbofan engine will
consume about 2% extra fuel and run at about ~0F hotter
turbine temperature in order to provide aircraft system
bleed air. The present invention addresses these problems
and deficiencies characteristic of the prior art and
conventional apparatus used to supply aircraft bl~ed air.
Figure 1 schematically illustrates an environmenta3
control system (ECS) typical of the prior art having a
conventional compressor bleed supply sy~tem 10 which
extracts compressor bleed air from an aircraft propulsive
gas turbine engine compressor sec~ion 8 to flow to and power
a conventional ECS pack 30, whlch is depicted using an air
cycle refrigeration syste~ to cool and condi~ion compressor
bleed air, as is typical of the prior art. Compre~sor bleed
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supply system 10 includes a compressor mid-stage air bleed
port 11 and a compressor discharge air bleed port 12 for
supplyin~ compressor bleed air through a compressor bleed
air line 9. Bleed air normally flows through mid-stage
bleed check ~alve 13 to shut-off valve 14. At low engine
power, discharge bleed valve 15 can be vpened, causing check
valve 13 to close and bleed air to be delivered from
compressor discharge port 12 to shut-off valve 14. Bleed
air pressure is reduced to a duct structurally safe level by
pressure regulator valve 16. Bleed air passes through
en~ine bleed air precooler 17, which cools compressor bleed
air, using cooler fan air ~rom engine fan 19, to a safe
temperature level before passing through duct 18 which is
typically located in the aircraft wing near the aircraft
fuel tanks. F~n cooling air from engine fan 19 flows
through a precooler temperature control valve 21 upstream of
precooler 17 and downstream of precooler 17 fan air is then
flowed overboard of the engine as indicated by arrow 17a.
Havinq been partially reduced in both pressure and
temperature, engine bleed air passes throuqh duct 18 to ECS
pack 30 for further temperature and pressure adjustment
before introduction to aircraft 50 including its cabin,
cockpit, and cargo bays as required. ECS 10w control valve
31 drops bleed pressure substantially so that the pressure
losses across ECS pack 30 maintains the desired bleed flow
from the engine to the aircraft. The pressure losses across
ECS pack 30 and in particular flow control valve 31 are very
expensive in terms of fuel and thrust because of the energy
spent compressing the bleed air.
A portion of the bleed flow passe~ through an ECS
refrigeration compressor 32 to a ram air heat exchanger 33
then to an ECS refrigeratlon turbine 34. ECS re~rigeration
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compressor 32 and ECS refrigeration turbine 34 are supplied
as a single ECS turbocompressor assembly wherein ECS
refrigeration compressor 32 is a centrifugal compressor and
ECS refrigeration turbine 34 is a radial inflow ~urbine.
The pressure drop across ECS refrigeration turbine 34 causes
it to drive ECS refrigeration compressor 32.
Heat is removed from the bleed air ~y heat exchanger 33
which receives cooling air from a conventional ram air scoop
35a located on an ou~side surface of the aircraft and
controlled by a ram air door 35. ~am cooling air is then
dumped overboard as indicated by arrow 33a. Other portions
of the bleed air bypasses ECS refrigeration compressor 32
through compressor bypass valve 36 or ECS pack bypass 37.
ECS pack bypass air from valve 37 mixes with bleed air from
ECS refrigeration turbine 3~ for ~inal temperature control
of conditioned ECS bleed supply air that is supplied to
aircraft 50 through an ECS air supply line 54.
Conditioned bleed air then flows to aircraft 50 for
passenger flesh air consumption, cabin pressurization and
temperature control. Temperature control i5 re~uired to
counter a varying cabin heat load 51 that includes aircraft
skin cooling, solar heating, passenger body heating and
electrical load heating. Part of cabin supply air 52
returns through re-circulation fans 53 and mixes with
~5 conditioned ECS bleed supply air from ECS air supply line
54. Cabin pressure is controlled by an outflow valve 55
that dumps overboard, as indicated by arrow 55a. Air flow
from outflow valve 55 is essentially compressor bleed air
taXen from engine compressor section 8 and supplied to the
aircraft.
Another problem addressed by the present invention
relates to aerodynamic drag associated with engine nacelles,
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wings, pylons, tail sections and other aircraft outer
surfaces. As air flows on to and over a surface such as an
engine nacelle and aircraft wing it progressively builds up
a low velocity boundary layer of increasing thickness.
S Within this boundary later a portion of the velocity
component of free stream total pressure is converted to
increased static pressure. As the result of rise in static
pressure, boundary layer thicXness, and diffusion a point is
reached where back pressure causes an otherwise laminar
boundary layer to become turbulent. In the turbulent
region, a considerable amount of total pres~ure is converted
to static temperature represented thermodynamically as an
increase in entropy. By the time the boundary layer laaves
the surface, or in the particular case of an aircraft gas
turbine engine, the end of the nacelle, an unrecoverable
loss in total pressure has occurred. The large entropy rise
associated with turbulence is at the expense of air
momentum. Turbulence also gives rise to increased static
pressure which may increase the intensity of rearward acting
pressure force on the surface. Now, if the boundary layer
thickness is kept small, separation and turbulence will not
occur or will be delayed and drag can be substantially
reduced.
It is well known that one way to avoid increases or to
reduce the build up in boundary thickness is to pump or
bleed of~ boundary layer air through holes in the aircraft
wetted surfaces such as the wing, tail, or portions of the
engine nacelle. Boundary layer pumps or compressors would
be desirable from an aerodynamic standpoint but, because of
the relatively large airflow rates and added weight and
complexity associated with effective bounda~y layer pumpinq
or bleeding, the concept has not been adapted in modern
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aircraft and engines. Therefore, in one embodiment of the
invention, this invention pro~ides a means for effectively
and economically using engine compressor bled air to power a
nacelle boundary layer ~leed compressor to bleed off laminar
flow boundary layer air from the nacelle to reduce drag.
A similar problem was addressed in, and reference may
be made to, patent application S/N 07/531,718, invented by
George A. Coffinberry, filed June 1, 1990, and assigned to
the same assignee and incorporated herein by reference.
Mechanically powered means for reducing boundary lay~r
drag of various aircraft parts ~uch as wings, nacelles, and
aircraft tail assemblies have been proposed in the past and
in patent application ser. No. 07/489,1SO entitled "AIRCRAFT
ENGINE STARTER INTEGRATEP BOUNDARY BLEED SYSTEM~, invented
by Samuel Davison, filed March 6, 1990 and assigned to the
same assignee. Other inventions addressing this problem are
disclosed in a patent application Ser. No. 07/531,718
entitled "GAS TURBINE ENGINE POWERED AIRCRAFT ENVIRONMENTA~
CONTROL SYSTEM AND BOUNDARY LAYER BLEED", filed on June 1l
1990, and a patent application Ser. No~ 07/572,825 enti~led
"AIRCRAFT ENGINE ELECTRICALLY POwERED 80UNDARY LAYER BLEED
SYSTEM ", filed on June 1, 1990, both invented by the same
inventor of this patent and assigned to the same assignee,
both incorporated herein by reference.
Another patent application Ser~ No. 07/531,734 entitled
l'GAS T~BINE ENGINE FAN DUCT BASE PRESSURE DRAG REDUCTION",
invented by the same inventor of this patent, filed on June
1, 1990, and assigned to the same assignee, and incorporated
herein by reference, discloses a mechanically powered means
of bleeding boundary layer air and reducing the drag of the
aircraft by introducing at least a portion of the
pressurized bleed air into the fan duct of the engine to
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reduce the base drag of the duct.
6UMMARY OF T~ IN~ENTION
Briefly, in accordance with one aspect of the present
invention, an environmental control system for supplying
aircraft or cabin air to an aircraft powered by a gas
turbine engine is provided with a propulsive engine
compressor bleed supply means and an ECS air ~low control
means which includes an energy recovery means for returning
the unrequired or unused amount o~ energy back to the
engine. The ECS flow control and energy recovery means
includes an air turbine powered by bleed air from the
compressor bleed supply means and is mechanically linked in
driving relationship to an auxiliary compressor that pumps
boundary layer air from the surface of the engine nacelle
and exhausts this air rearward of the en~ine to produce
thrust.
The air turbine includes a variable inlet nozzle and a
means to control it which is in switchable ~luid supply
communication with sources of gas turbine engine compressor
bleed air and provided with a means for ducting its exhaust
flow to the environmental control system ECS pack. The ECS
pack includes a means ~or producing a signal that is
provided to a variable inlet control means for controlling
the variable inlet turbine nozzle.
ADVANTAGE8
Among the advantages provided by the present invention
is the ability to eliminate waste of ~uel and thrust such as
tha~ incurred in dropping the pressure of propul~iYe engine
bleed aix by ECS pack flow control valves. The present
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inventio~ provides a more efficient aircraf~ gas turbine
engine ~y using thP energy in the compressor bleed air that
is conventionally wasted.
Another advantage of the present invention is that it
allows the use of a lighter weight and more efficient and
longer life aircraft gas turbine engine.
The present inventionls use oP boundary layer air to
precool compressor bleed air has the ~urther advantag~ of
eliminating a need for a drag producing fan cooling aix
scoop and reduces the size and weight required for the
precooler.
By using the otherwise wasted energy of compressor
bleed air to drive the boundary layer suction compressor the
present invention is able to reduce engine nacelle boundary
layer drag in a fuel efficient manner. Furthermore, this
advantage can be optimized because boundary layer dra~ is a
penalty that accumulates over the cruise portion of the
flight envelope and is therefore an easy point to design
for.
These advantages and other features and advantages will
become more readily apparent in the following description
when taken in conjunction wi~h the appended drawings.
BRIEF DESCRIPTION OF ~ D~WIN~
The ~oregoing aspects and other fsatures of the
invention are explained in the following description, taken
in connection with the accompanying drawin~ where:
FIG. 1 is a schema~ic representation of a prior art ECS
system for supplying engine bleed air to an ECS pack and
conditioning it for use in an aircra~t.
FIGo 2 is a schematic representation of an aircraft
propulsive gas turbine engine boundary layer bleed syste~
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for an aircraft propulsive gas turbine engine powered
environmental control system (~CS) in accordance with the
preferred embodiment of the present invention.
FIG. 3 is a schematic representation of an energy
conserving ECS pack for use with the ECS of FIG. 2 in
accordance with the preferred embodiment of the present
invention.
DETAILED D~SCRIPTION OF THE INVEN~ION
The present invention provides an ECS having an energy
recovering compressor bleed flow control means 40 as
illustrated in FIG. 2. Schematically illustrated in FIG. 2
is compressor bleed flow control means 40 used to control
the flowrate of compresor bleed flow through duct 18 in
FIG. 2 to an energy conserving ECS pack 80 of FIG. 3 which
is in accordance with the preferred embodiment of the
present invention.
FIG. 2 schematically illustrates a part of an ECS
including a compressor bleed supply means 59 that is
partially controlled by compressor bleed flow control means
40 to supply compressor bleed air to energy conserving ECS
pack 80 in FIG. 3 through duct 18. Compressor bleed supply
means 59 includes a mid-stage bleed port ~1 and a compressor
discharge bleed port 62 that bleed compressed air from
corre~ponding stations in the comprecsor sections of an
aircraft propulsive gas turbine engine 60 and are controlled
by mid-stage and compressor discharge bleed control valves
61A and 62A respectively. Compressor bleed supply means 59
supplies compressor bleed air to compressor bleed flow
control means 40 that also serves as an energy recovery
means that uses a portion of an unrequired or unused amount
of energy to reduce boundary layer drag and return another
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portion of this unused amount of energy back to engine 60 in
form of thrust.
The energy recov~ry means includes an air turbine 63
drivenly connected through a shaft to a boundary layer bleed
compressor 64. Air turbine 63 is preferably a single stage
radial inflow turbine having a variable area turbine inlet
nozzle 65. Turbine nozzle actuation means 66 receives a
signal 83 from aircraft energy conserving ECS pack energy
conserving 80 of FIG. 3 and opens and closes its exhaust
area so as to vary and control the amount of ble~d air taken
rom either mid-stage bleed port 61 or compressor discharge
bleed port 62. Variable area turbine nozzle 65 converts a
portion of it~ supply pressure to high velocity air
appropriately directed on to the inlet of air turbine 63.
Air turbine 63 converts this high velocity air to power
which drives boundary layer bleed compressor 64.
The drop in pressure from engine bleed port 61 or 62 is
thus converted to compressor drive energy while at the same
time the amount of bleed air taken from the engine is
controlled by varying turbine nozzle area.
In a conventional prior art bleed air arrangement, as shown
in FIG. 1 and described previously, comparable pressure drop
would be simply throttled by EC5 flow control valve 31 in
FIG. 1 and energy would be lost.
Boundary layer bleed compressor 64 pumps air through
engine nacelle suction holes 67 from an aircraft wetted
surface represented in the preferred embodim~nt by a nacelle
surface 68. Removal of boundary layer air through suction
holas 67 prevents build up of a thick region of ~oundary
layer stagnation and transition to turbulence, thu~
maintaining laminar flow over nacelle surface 68 and
reduction in nacelle surface frictional drag. Reduction in
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drag permits engine 60 to be operated at a lower level of
thrust and therefore burn less fuel.
Boundary layer suction air from compressor 64 flows to
a precooler heat exchanger 69 where it cools and in return
is heated by engine bleed air leaving turbine 63 for
delivery to enerqy conserving ECs pack 80 in FIG. 3 through
duct 13 for eventual use by aircraft 50 as what is commonly
referred to as cabin air.
Higher temperature boundary layer suction air then
flows to thrust nozzle 71 where it exhausts at a velocity
the same or slightly higher than aircraft flight velocity.
So long as the amount of suction air does not exceed an
amount which would cause an increase in free stream
penetration into the stagnation region associated with the
laminar boundary layer, the ram drag penalty, due to change
in momentum of the boundary layer air that is suctioned off,
is essentially 2ero. Therefore, essentially all of the
momentum associated with air leaving nozzle 71 is available
as positive thrust. Therefore, energy derived from turbine
63 is returned to the engine in the form of both drag
reduction and thrust increase.
For conditions in which the demand of bleed air exceeds
the capacity of noz~le 65 or turbine 63, bleed air may
bypass turbine 63 through a bypass valve 72. Overpressure
valve 73 provides overpressure protection and shutoff as in
the case o~ a conventional system.
Precooler heat exchanger 69 replaces fan precooler 17
o~ a conventional prior art ECS system (shown in FIG. 1) and
is relatively small since compressor 64 requires less
pressure ratio than the engine fan and therefore results in
colder air into the cold side o~ the heat exchanger~
Turbine 63 cools engino bleed air by work extraction before
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the bleed air reaches heat exchanger 69. After leaving heat
exchanger 69, the bleed air is substantially colder than it
would be for a conventional prior art ECS system using
precooler 17 shown in FIG. 1.
FIG. 3 schematically illustrates an eneryy conserving
ECS pack 80 in accordance with the preferred embodiment of
the present invention for use wi.th energy recovering
compressor bleed flow control means 40 as described above
with the previous discussion. A flow sensor 81 is used to
measure the amount of bleed air flowed from tha engine
compressor through duct 18. A controller 82 receives a
flowrate signal through flow rate signal lines 81a and then
sends an appropriate control signal 83 to turbine nozzle
area actuator 66, as shown in FIG. 2, in response to a
demand signal 84 received from the aircraft and the flowrate
signal.
In order to reduce the bleed air supply pressure to the
ECS pack and thereby provide more power to turbine 63 of
FIG. 2, bypass valve 37 is opened to a maximum extent, thus
reducing the amount of flow and associated pressure drop
through the rest of energy conserving ECS pack 80. This
extra degree of bypass is made possible by the fact that the
invention described previously by FIG. 2 provides colder air
to energy conserving ECS pack 80.
The remainder of ECS pack bleed air flows either
through ECS refrigeration compressor 32 or compressor bypass
valve 36. ECS refrigeration compressor 32 must be bypassed
through valve 36 when there i5 insufficient pressure drop
across ECS re~rigera~ion turbine 34 to powex ECS
refrigeration compressor 32.
Ram air door 35 opens to provide ram ~ir cooling air to
heat exchanger 33 and proper temperature con~rol o~
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compressor bleed air passing through ECS bypass valve 37,
compressor bypass valve 36, and ECS refrigeration compressor
32. The compressor bleed air then passes through ECS air
supply line 54 to aircraft 50 (as shown in FIG. 1).
Components and arrangement of eneryy conserving ECS pack 80
that are essentially the same as those in the conventional
ECS system illustrated as prior art in FIG. 1 are labeled
with the same numerals.
The present invention provides a distinct advanta~e for
improving specific fuel consumption by returning otherwise
unused energy to the en~ine and using a variable area inlst
nozzle turbine to drive a compressor for purposes of laminar
flow suction~ engine bleed air cooling and additional thrust
from the suction air. An example of the magnitude of the
advantage that can be realized from the present invention
has been calculated based on the following typical aircraft
configura~ion and operatin~ conditions. A McDonnell-Douglas
MD-ll aircraft having two Gcneral Electric CF6--~OC2 engines,
a 280-passenger load, and operating at a 35,000 foot
cruising altitude at a point where each engine is
conventionally operating at a thrust level o~ 7919 pounds.
It ha~ been calculated that for this typical configuration
and these operating conditions that each of the two enqines
should have a 1.3% improvement in engine specific fuel
consumption (SFC).
The present invention has been described in an
illustrative manner. It is understood that the terminology
which has been used is intended to be in the nature of words
of description rather than of limitation. Unused engine
bleed air energy in the form of pressure and temperature can
be used to power a turbine and provide bleed airflow
control, bleed air temperature reduction, and use~ul power
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such as to drive a laminar flow suction air compressor. The
same unused energy might also power an air ejector for
suction flow or for other reason Turbine power might be
used to drive a motor-driven compressor for the same
advantages of the present invention or for other reasons.
Obviously, these and many other modifications and variations
of the present invention are possible in light o~ the above
teachings. It is, therefore, to be understood that within
the scope of the appended claims, the present invention may
be practiced otherwise than as specifically described.
